Crewed Mission to Mars--Jan. 2018

Compare to TOPAZ-II, which I believe is one of the most developed light-weight designs; that would generate 5 kWe at a weight of almost one metric ton.

Topaz-2 used a thermoelectric conversion mechanism, which has extremely low efficiency. That 5-6.6kWe was derived from 150kW thermal. Anything even resembling a turbine will be far, far better.

However, it will also be heavier (the weight of the turbine conversion mechanism) and less mechanically reliable (the last is an important point, since these are usually supposed to run for some time with no maintenance). Even the craziest lightweight reactor designs you can point me at can't compare with the power-to-weight ratios of flown solar cell designs--not just IKAROS, but Ultraflex arrays have power-to-weight ratios of several hundred watts per kilogram.

How long do nuke subs stay submerged? Best I can find is a mention of 180days for the earlier nuke boats, several of the problems mentioned here have already been (partially) solved.

Food amounts I see are being calculated on earth norms, without gravity yanking us down and putting stresses on our joints, do we really need as much energy (2kCal?) from our foods? The ISS isnt quite the same as interplanetary travel, but it would be interesting to see dietary changes, Im sure calcium suppliments to retard bone density loss is a given.

Silly question (mebbe I should put it my own thread, ah well) - I know there are various perigrees and orbital slingshots that work, earth to lunar orbit. Is it 'easier' to do Earth to Mars or Moon to Mars? I know we dont have a lunar base (officially ) yet, but with a significantly lower gravity (1/6th?), does that correspond to a similarly lower escape velocity / delta vee required - making lunar launches cheaper per tonne?

With no air resistance and lower gravity is a lunar 'railgun' style launch on the table?

Oh and lastly, if they do send a ship, it best be called the Armstrong.

Nuke subs are immersed in water so this helps with a few things (heat exchange, onboard oxygen generators) they can also surface if shit goes bad. They also have nuke plants so power isn't a problem and they built and launched in dry docks so can weight a lot without a problem.

Crewing a nuke or attack sub is very demanding, so they eat very well. Food is the limiting factor on sub deploys. They could probably stay down a year if they only ate MREs. Unless they killed each other first.

Compare to TOPAZ-II, which I believe is one of the most developed light-weight designs; that would generate 5 kWe at a weight of almost one metric ton.

Topaz-2 used a thermoelectric conversion mechanism, which has extremely low efficiency. That 5-6.6kWe was derived from 150kW thermal. Anything even resembling a turbine will be far, far better.

However, it will also be heavier (the weight of the turbine conversion mechanism) and less mechanically reliable (the last is an important point, since these are usually supposed to run for some time with no maintenance). Even the craziest lightweight reactor designs you can point me at can't compare with the power-to-weight ratios of flown solar cell designs--not just IKAROS, but Ultraflex arrays have power-to-weight ratios of several hundred watts per kilogram.

From my armchair perspective, however, why is it important to talk about power to weight ratios without talking about how much thrust and/or specific impulse you can achieve with that power? An incredibly efficient, light-weight solar design that simply can't deliver enough juice to run future engine concepts (MPDT, VASIMIR, etc.) doesn't seem very useful if you want to get anywhere quickly.

Perhaps I'm not thinking this through correctly (hell, that's likely), but it seems spending 25K Kg for a ComStar reactor producing 15 MW would be more than made up for by hanging 10 200 Newton MPD thrusters off of it with an exhaust velocity of over 100 KM/sec. Assuming your spacecraft weighed in at 80,000 Kg, what would be your speed after 90 days at 2000 newtons of thrust?

Yes, this is all very, very speculative tech... Gen4 hasn't built a ComStar reactor yet due to licensing timelines, and 200 N MPDTs are years away. Tito isn't going to get a spacecraft like that in 5 years... I'd think more like 20-25. But I would think putting money into power supplies and high-energy propulsion would be a better use of his cash.

Perhaps I'm not thinking this through correctly (hell, that's likely), but it seems spending 25K Kg for a ComStar reactor producing 15 MW would be more than made up for by hanging 10 200 Newton MPD thrusters off of it with an exhaust velocity of over 100 KM/sec. Assuming your spacecraft weighed in at 80,000 Kg, what would be your speed after 90 days at 2000 newtons of thrust?

That is a very small force, but applied constantly for 90 days (with assumptions of no drag, no obstacles big or small and no gravimetric interferences) then the space craft would be going approximately 8,100 m/s (about 0.00003c).

From my armchair perspective, however, why is it important to talk about power to weight ratios without talking about how much thrust and/or specific impulse you can achieve with that power?

Because the power to weight ratio is one major factor determining the mass that your thruster needs to push at any given time; the major factor with electric propulsion spacecraft, because the life support/payload is essentially fixed by the mission and the specific impulses on even existing electric thrusters are high enough that you have quite low propellant requirements. Actually, increasing specific impulse too much can increase your needed mass, because specific impulse depends on the energy you’re putting into the propellant; at a certain point, the necessary increases in mass for your power system outweigh the savings in propellant. I believe the optimum figure for Mars is somewhere in the general range of 3000-4000s, which has been achieved by existing thruster designs.

Er, try thinking this through a little more. If you need more power...add more solar panels. That’s another advantage of solar over nuclear; it’s by its nature much more flexible in terms of total power output, whereas nuclear systems tend to be more monolithic. Handy for a step-by-step buildup.

NervousEnergy wrote:

Perhaps I'm not thinking this through correctly (hell, that's likely), but it seems spending 25K Kg for a ComStar reactor producing 15 MW would be more than made up for by hanging 10 200 Newton MPD thrusters off of it with an exhaust velocity of over 100 KM/sec.

So, you have a 25 mT nuclear reactor that can produce 15 MW of power, or about 600 W/kg. That’s nice...about as good as Ultraflex panels. However, from a (very brief) look around, that mass does not include things such as structure, cooling, power conversion, and the like, so you’re likely to be spending rather more than that; I hesitate to guess a number because I am neither a nuclear nor an aerospace engineer, let alone both, and while I do have some documentation on conceptual designs for NEP and SEP spacecraft most of it dates from SEI and so is technically rather obsolete.

In any case, compare that with the situation where you spend a similar amount of mass on solar cells. In the case of Ultraflex-type cells, you would get a comparable amount of power for the mass (although the solar cells would need more structure to support them than the nuclear reactor, the total mass would likely be comparable to the structure + radiator needed by the nuclear plant), but, rather advantageously, you don’t need to develop any new technology or space-qualify them or anything of that sort, since they exist today and are already designed for space use. ATK would have to scale up production and some redesign would be needed because the circular profile arrays probably aren’t workable for extremely high energy applications like that, but overall it would be a lot less work.

If you instead used IKAROS-performance cells...well. Remember, those produce 1000 W per kilogram; that is, 25,000 kg of them will produce 25 MW--a 66% increase in power over your nuclear reactor. You may be able to get away with a rather lightweight structure if you design things properly as well, since these are literally solar cells made out of ultrathin-film solar cells. This would require a lot more R&D to scale up to the necessary manufacturing quantities and to develop techniques for orbital assembly than Ultraflex-type arrays, but on the other hand...I’ve seen people whom I think are reasonably intelligent and trustworthy say that they might be able to hit 3-5 times that performance in power-to-mass ratio. In other words, produce 75-125 MW of power for the same 25,000 kg in cells. Naturally, quite a lot of mass will be needed in power conversion devices and structure in that case--while the cells might be outputting that much energy, it’s not like it just goes straight to the thrusters in the proper input...format? might be the best word (by which I mean, the cells will obviously output DC power, whereas the thrusters might need AC 3-phase, say). That will bring down the effective power-to-mass ratio...how much, hard to say. The large size of the arrays will also require a certain degree of delicacy in maneuvers to avoid damaging them through overacceleration.

Still, you’re getting perhaps 5-8 times as much power for the same weight, or using less weight to get the same amount of base power, so...yeah.

NervousEnergy wrote:

Assuming your spacecraft weighed in at 80,000 Kg, what would be your speed after 90 days at 2000 newtons of thrust?

That is...really, really bad mission design and the complete wrong question to be asking. For one thing, to actually get anywhere with constant acceleration, you need to follow a brachistochrone trajectory, which involves speeding up and then slowing down. Otherwise, you’ll just blow right past Mars in a hyperbolic orbit. Although using a Mars Flyby Rendezvous mode has been proposed a number of times in the past to minimize mission mass, this is not really a form of that. If you accelerate at a certain rate a for part of a trip and decelerate at the same rate (just in the opposite direction) for another part, then

d = 2a*t_f*t_m - 1/2*a*t_f^2 - a*t_m^2

where d is the total distance traveled, t_f is the total amount of time elapsed, and t_m is how much time elapsed between when you started accelerating and when you started decelerating. This assumes you are going from “rest to rest,” which is decidedly NOT true for interplanetary travel, but which is acceptable for a first cut analysis. For the moment using your numbers...t_f is 90 days, or 7,776,000 seconds. a is 2000 N/80,000 kg = 1/40 m/s^2. t_m is a bit hazy, but take it to be 45 days, or 3,888,000 seconds. Then, your hypothetical space ship will travel 377,913,600 km in 90 days, about the diameter of Earth’s orbit around the Sun. Very impressive so far, but...

It takes propellant to speed up and slow down. This was, obviously, not accounted for at all in the previous analysis (a was constant). Fortunately, it’s pretty easy to compute how much would be needed provided you throttled to maintain a constant thrust. Over the 45 days of initial acceleration, a delta-V of 97,200 m/s is applied. The same delta-V is then applied during the deceleration to slow the spacecraft to a stop. The Tsiolkovsky rocket equation states that

m_r = exp(delta-V/v_e)

where m_r is the mass ratio (initial mass divided by final mass) and v_e is the exhaust velocity. Continuing to use your provided numbers, v_e is 100,000 m/s, so m_r is about 2.64. This needs to be applied twice; first to the final spacecraft mass, then to the spacecraft mass at turnaround, so the overall mass ratio is 6.9696, ie. 7 more or less. So if your spacecraft masses 80,000 kg at the end of its mission, then it masses 560,000 kg at the beginning; or if it masses 80,000 kg at the beginning, then it can only mass 11,428 kg at the end. A bit of a problem there, given that you specify that the reactor masses 25,000 kg by itself...

Of course, those numbers feed back into each other. For instance, the spacecraft will obviously not accelerate as quickly if it masses 560,000 kg than if it masses 80,000 kg, so it will experience less of a delta-V (and won’t travel as far), so the mass will be a bit less, so it will accelerate a little bit more quickly, so it will experience a greater delta-V, etc., so these mass numbers are, strictly speaking, wrong. However, I think they do rather usefully illustrate the magnitude of the problem and also begin to give a considerable insight into the next issue with your “concept,” which is that electrically powered missions are very difficult to analyze, especially for timelines, to the point where I will not even try (if you gave me a couple of months and SPICE, I could probably whip something up, but short of that). The thrust schedule is a lot more complicated than “turn them on and orient the spacecraft prograde”. This makes questions like “How fast would it be going after 90 days of thrust?” not very helpful.

Finally, you’re also falling into a significant mission analysis and design trap that I like to call Optimizing for the Wrong Thing. Basically, it happens when someone sees a problem and a possible solution to the problem, then tries to whang away at the solution given more and more constraints until it fits the problem, without considering whether there might not be a better solution to the original problem. For a real-world example of this happening, consider the space shuttle. The problem, of course, was high launch costs. The solution, everyone thought at the time, was building a reusable spacecraft. So a lot of rather smart people at NASA, Grumman, Lockheed, North American, and so on worked very very hard to design a reusable spacecraft, even as additional political constraints (eg., a cap on total development funding) and technical constraints appeared. Eventually, they did design a spacecraft which was mostly reusable, but which didn’t actually save any money because of all the compromises that were needed to build it under those constraints. For instance, instead of having reusable flyback boosters, they used solid rocket boosters that could be recovered and refurbished because those would be cheaper to develop and supposedly nearly as cheap operationally. It turned out that shipping SRB segments all the way back to Utah for recasting was not cheaper than simply building new ones, and moreover segmenting the solids (necessary for shipping them) introduced a weak point which failed during Challenger and killed seven astronauts.

Essentially, what you’ve done is see a problem: Spending a Long Time in Space is Bad for You. You’ve decided on a solution--Don’t Spend a Long Time in Space. And now you’re searching for ways to make that happen using (mostly) nuclear-electric propulsion, without considering possible alternative approaches or even whether going to Mars for short periods of time is worthwhile to begin with (a lot of people would say no, a 20-30 day stay on Mars isn’t worth it). For instance, we’ve gotten better at mitigating the effects of microgravity on human bodies over the last twenty years, with better exercise routines. If, instead, the astronauts were in partial-gs, then it might (we don’t know, having never had organisms spend significant amounts of time in fractional-g environments) be enough to basically remove the degrading effects of microgravity altogether, especially combined with possible drug interventions (eg., anti-osteoporosis drugs). So instead of developing systems to allow a sub-year round trip to Mars, you would design a centrifuge or artificial gravity. Much easier in some ways. Similarly, one of the biggest advantages of electric propulsion over chemical propulsion is that it requires much less propellant mass because of the far greater ISP of electric thrusters compared to chemical thrusters. You could use some of this saved mass to improve passive radiation shielding on the habitat. You could also use excess power from your power source to power an EM shield. Although it might not be capable of decreasing flux to Earth-surface levels, it could improve protection over passive shielding for less mass per milliSievert reduction in dose. You could send older astronauts, who would experience less risk from radiation doses than younger ones (baldly, they’re more likely to die before they can develop cancer or heart disease or whatever). And so on and so forth. There are a lot of ways other than Don’t Spend a Long Time in Space to mitigate Spending a Long Time in Space is Bad for You, and they may be more attractive given present and near-future technology levels, or overall.

Compare to TOPAZ-II, which I believe is one of the most developed light-weight designs; that would generate 5 kWe at a weight of almost one metric ton.

Topaz-2 used a thermoelectric conversion mechanism, which has extremely low efficiency. That 5-6.6kWe was derived from 150kW thermal. Anything even resembling a turbine will be far, far better.

However, it will also be heavier (the weight of the turbine conversion mechanism) and less mechanically reliable (the last is an important point, since these are usually supposed to run for some time with no maintenance). Even the craziest lightweight reactor designs you can point me at can't compare with the power-to-weight ratios of flown solar cell designs--not just IKAROS, but Ultraflex arrays have power-to-weight ratios of several hundred watts per kilogram.

The best nuclear reactors built achieved 430,000W/kg (thermal) back in 1967.

Granted they weren't built for longevity but if we're talking about the state of the art then that's what it is.

For instance, instead of having reusable flyback boosters, they used solid rocket boosters that could be recovered and refurbished because those would be cheaper to develop and supposedly nearly as cheap operationally.

Was there the technical or engineering capabilities to build a flyback capable booster in the 70s when the shuttle was being designed? If not, the the line of reasoning is moot.

The simple reality seems to me to be that we're only now really just starting to talk about boosters as something that isn't purely disposable. Even then, you've got what? SpaceX working on flyback capabilities for the Falcon9, the Russians showing a mockup of the Baikur booster in 2001 that never went into production AFAIK, and a USAF research project that ran from 2010 to 2012.

Thanks for that in-depth post, TiL. My insistence in looking at nuclear power sources came from an obviously incorrect assumption that it was impractical to use enough solar cells to provide the MW range power output required for high energy thrusters. I also didn't think the propellant requirements would be that high given the efficiencies. Looks like the thrusters need to get to considerably higher exhaust velocities for the power budget to make it practical to go gadding about the inner solar system at full acceleration / deceleration for the entire trip. Creating a complex vehicle with a centrifuge for the living hub (HAL, open the pod bay doors, please...) is a well-considered idea, but the engineering complexity increase is daunting.

The only thing I would quibble with in the post is discounting a '20-30' day stay as a poor mission design criteria. Landing and lifting human cargo alone is a significant expansion of the mission in terms of risk and cost considering all the unknowns. How many times did we loop around the moon, testing our ability to get there and back, before attempting a landing? Now it could be argued that the cost/risk of going all the way to Mars makes it more viable to go for the gold and do it all in one trip, but I can see the arguments lining up against it. And even if we did, 20-30 days is a LONG time on the surface... successfully putting a crew down for that long and getting them home would be seen as an almost unbelievable achievement. I'm skeptical there's just a ton of science that can be done on the surface that can't be done in 20 days that we're not doing now, with multiple high-capability robots wandering about the planet. Just landing long enough to dig up a bunch of surface material and actually bring it back would be huge.

From what I've been able to find/read about the state of the art in high-energy thrusters, MPDT units have the highest theoretical performance (110K/sec exhaust for a 1MW power budget at peak efficiency giving 200 N of thrust using Hydrogen propellant). My rough example above assumed a much higher energy need for that engine. Now I'm wondering how little weight one can get away with for the energy and propulsion systems while minimizing trip time. I can't find any estimated weight values for the engines themselves.

We don't have to get to Mars in a few weeks for the trip to be acceptable, obviously, but even with your non-velocity related 'space is bad for you' mitigation efforts, I still think that psychologically we need to get the round trip under 1 year, and preferably shorter.

Was there the technical or engineering capabilities to build a flyback capable booster in the 70s when the shuttle was being designed? If not, the the line of reasoning is moot.

The simple answer is yes, the capabilities existed at the time.

The longer and more complex answer is yes, they could have built flyback boosters, but there would have been issues and they, too, might not have been cheaper than disposable boosters.

The really long and complex answer starts from the beginning of time (aka 1945), when von Braun started really thinking about spaceflight because of a certain...forced inactivity following his departure to the United States from Germany. He really wanted to launch a Mars mission, and started working on numbers to show that such a mission would be practical, using then-modern technology. The resulting mission was REALLY BIG in modern terms, with something like 37,000 metric tons of spacecraft, propellant, astronauts, equipment, and so on. needing to be put into low Earth orbit; that is, it had an "IMLEO," "Initial Mass in low Earth orbit," the usual figure of merit, of 37,000. The abortive Space Exploration Initiative of Bush Sr. was widely castigated as having a monstrously large Mars spacecraft, which would have had an IMLEO of at maximum about 1,000, and only for the worst opportunities at that (because of celestial mechanics, missions to Mars get more or less difficult at certain times).

Obviously, he didn't consider it practical to build a rocket capable of lifting 37,000 mT of payload into space in a single launch. In fact, he didn't consider it especially practical and economical to build a rocket capable of lifting a mere thousand tons, or even a hundred tons into space in a single launch. Instead, he figured on a rocket capable of carrying around 25 metric tons of payload into space per launch (sound familiar...? That's about what the Shuttle was capable of, although for very different reasons). To launch so much payload in a reasonably short period of time and for a reasonable amount of money, these absolutely had to be reusable, fully reusable as a matter of fact; he figured on 46 vehicles conducting 950 launches (!) over an 8 month period (!!--that's about 4 launches per day!) to build his mission. Building 950 vehicles in eight months would simply not be practical at all, then or now.

And so the idea of a shuttle was born. It didn't exactly stay there, though; not only was von Braun possibly one of the greatest space popularizers of all time (Carl Sagan alone rivaling him, although in a very different way), but of course he gained a fair bit of influence over the American aerospace community, which began to generally feel that the logical order of development went shuttle-station-Moon-Mars (this is deeply pervasive even today). Moreover, for rather different reasons the Air Force in particular was considering lifting spacecraft for future development, at first (in the Brass Bell, BoMi, and RoBo projects) as a logical extension of the trend towards faster, higher-flying aircraft, then (in the Aerospaceplane and other concepts) as a goal unto itself (unfortunately, I don't have my best source on me just at the moment, so I can't go into great depth). As a matter of course, these would have been reusable and generally resembled the later X-30/NASP of the 1980s in concept. Even during the contractor competitions which led into Mercury and Apollo (not Gemini, which was developed originally as a straightforward upgrade to the former), the interest in lifting rather than capsule designs was widespread, with many of the proposals being such lifting spacecraft (capsules were chosen due to concerns about development time). Despite the Apollo Program, studies, especially again at the Air Force, continued at a low level throughout the 1960s.

So, when the Apollo Program started to wind down in about 1969 (and the really ambitious follow on got definitively killed in about 1970), it was pretty natural for everyone involved to go, "hm, well...shuttle is the next logical step". And they lobbied and maybe got too caught up in their own propaganda and Nixon wanted to get support in California for his reelection bid, and it ended up happening. But, as a matter of course, their shuttle would be fully reusable, just like almost every design since von Braun's first (excepting, in particular, one from Lockheed called Starclipper...but that's another story). There were a large variety of ideas for how this might work, but typically the vehicle which would actually go into space would take off connected to another, much larger vehicle called the booster. Usually, these would share engines and overall design, but the booster would be much larger...like, empty it would weigh as much as a loaded 747. But there were exceptions; the Trimaran concept, for instance, had three identical vehicles taking off together, somewhat similar to the Delta IV Heavy. One would go into space, the other two would just be boosters, but any vehicle could take either role.

It turned out that developing the booster would cost a lot of money, about as much as the orbiter itself (at least according to projections). Because Congress and the OMB (ie., the Nixon administration) didn't want to give them that much money, they redesigned the Shuttle to use an external tank (instead of carrying propellant internally) and to use solid rocket boosters similar to those used on the Titan IIIC instead of the big booster. And that's how the Shuttle was made.

However. If you look at that brief description of the booster I gave earlier, you'll see that it was supposed to be pretty similar to the orbiter in most cases...same shape, same engines. Clearly, especially in the latter case, those designs were workable; the Shuttles flew for thirty years, and the SSMEs were quite reliable subsystems. And because the booster never goes as fast as the orbiter, it suffers much less reentry stress and overall is exposed to a less extreme environment. There's no question at all that they could design it to survive reentry. The only real problem with them was the sheer bloody size of the thing; they were airplanes that had to swallow the whole external tank and more to function, and were just as big and heavy as that implies--the An-225 would be diminutive compared to these guys. But could they make it work...? Yeah. I think they probably could.

Even discounting that, they could have built proposed large liquid pressure-fed boosters instead, which would also probably have been cheaper to refurbish than the solids. The main problem with the latter is that casting the fuel matrix--ie., "refueling" the booster--costs exactly the same with a used as a new booster. But a used booster has the costs involved with recovery and refurbishment included, which are also rather expensive. Manufacturing the case isn't exactly cheap, but it's also not exceptionally expensive; they don't have a lot of fancy equipment in them to save money on. By contrast, with a liquid booster most of the cost is contained in the hardware, the complicated turbopumps and so on, and you theoretically save a lot of money by building once, repairing twice (so to speak). Fuel is essentially free, and you just pump it in.

NetMasterOC3 wrote:

The simple reality seems to me to be that we're only now really just starting to talk about boosters as something that isn't purely disposable. Even then, you've got what? SpaceX working on flyback capabilities for the Falcon9, the Russians showing a mockup of the Baikur booster in 2001 that never went into production AFAIK, and a USAF research project that ran from 2010 to 2012.

This mostly just shows (to put it bluntly) ignorance of the past. For instance, the flyback capabilities of the Falcon 9 are not really any different than the capabilities of the Kistler K-1, albeit the specifics are different (the K-1 was boostback and used parachute landing, but both aimed to be reusable and land at launch site). It's just that the K-1 was from the 1990s. Similarly, NASA conducted some fairly extensive studies for liquid flyback boosters (under that specific name) in the 1990s with Boeing and Lockheed Martin because of improvements to Shuttle safety and lift capability, but which never advanced much because of a lack of funding and eventually died after the Columbia disaster. During the 1980s, you not only had X-30/NASP, which was supposed to be a SSTO (among other things), but you had a number of upgrade studies for the Space Shuttle, both before and after Challenger which proposed liquid and reusable boosters. Reaching back farther, even in the 1970s some of those were proposed...and in the 1960s and before, shuttles were considered fully reusable as a matter of course, as I previously mentioned.

NervousEnergy wrote:

The only thing I would quibble with in the post is discounting a '20-30' day stay as a poor mission design criteria. Landing and lifting human cargo alone is a significant expansion of the mission in terms of risk and cost considering all the unknowns. How many times did we loop around the moon, testing our ability to get there and back, before attempting a landing? Now it could be argued that the cost/risk of going all the way to Mars makes it more viable to go for the gold and do it all in one trip, but I can see the arguments lining up against it.

The usual precursor mission nowadays is a Mars orbit mission, to visit Phobos and the like. This eliminates the major delta-V and mass cost that comes with actually landing on the surface, and of course allows testing of hardware. You don't really need to go to Mars to test the in-space hardware, you just need to spend a year or two in space (which is one reason why NEO missions have been proposed as an intermediate between Mars and the Moon).

Also, we only went to the Moon twice before we landed there...and one of those was basically a propaganda stunt that didn't really prove anything (both the rocket and the capsule had already been tested). Only Apollo 10 was really a precursor mission, and arguably that wasn't necessary either (Apollo 9 had already tested the LM).

NervousEnergy wrote:

And even if we did, 20-30 days is a LONG time on the surface... successfully putting a crew down for that long and getting them home would be seen as an almost unbelievable achievement. I'm skeptical there's just a ton of science that can be done on the surface that can't be done in 20 days that we're not doing now, with multiple high-capability robots wandering about the planet. Just landing long enough to dig up a bunch of surface material and actually bring it back would be huge.

If you just want surface material, you don't need to land at all. Fully robotic Mars sample return (which NASA is presently de facto working on with the 2020 Curiosity II lander; basically, it's MAX-C reborn cheaper) can be done a lot sooner than a Mars landing. And if you just go into orbit, you can carry a lot of partially robotic sample return landers along for the mass cost of a human lander capable of doing a 20-30 day mission.

20-30 days is also not really all that long in the greater scheme of things, and especially in modern planning. Even modern lunar returns plan on 14-day or so stays, minimum; it's hard to say that it makes sense to bust ass and spend several times as much mass to go to Mars and then spend barely any more time there than on the Moon, especially given usual assumptions about travel time (and what those travel times entail--Venus flybys). Given the sheer expense and difficulty that comes with putting people on Mars, it doesn't really make sense to plan a mission that prevents them, for example, from doing long traverses to visit interesting sites a long way from their base camp (which will exist, of course), or revisiting sites that prove more interesting than originally thought, or just having some freeform planning (because a 20-30 day mission is going to be planned almost as tightly as Apollo). About the only way it makes sense is if you've built up an infrastructure that allows 20-30 day missions repeatedly for little marginal cost, but doesn't really allow for longer efforts.

Was there the technical or engineering capabilities to build a flyback capable booster in the 70s when the shuttle was being designed? If not, the the line of reasoning is moot.

The simple answer is yes, the capabilities existed at the time.

The longer and more complex answer is yes, they could have built flyback boosters, but there would have been issues and they, too, might not have been cheaper than disposable boosters.

It may have been technologically possible, but I think it would have been a lot harder back then, especially with propulsive landing.

It's not easy now and SpaceX got weirdly lucky. They developed an engine sized for pretty much the smallest viable launcher because they couldn't afford to do more. Which means they're stuck with a smaller engine than most launch providers would like and they have to use 9 of them for something big enough to be interesting. But one of those engines is small enough, just barely, to land a Falcon 9. An Atlas 5 couldn't do it, the one engine's too big. Even with a reusable Falcon 9, it will have to land with >1G of acceleration, it won't be able to hover or descend at a constant rate.

I think that was probably doable in the 90s, and we've done much harder things now like self driving cars, but it would have been extremely challenging in the 70s. A winged booster maybe, but that's less mass efficient. Or separate landing engines, but same problem.

It may have been technologically possible, but I think it would have been a lot harder back then, especially with propulsive landing.

Well, yes. It would have been.

Megalodon wrote:

It's not easy now and SpaceX got weirdly lucky. They developed an engine sized for pretty much the smallest viable launcher because they couldn't afford to do more. Which means they're stuck with a smaller engine than most launch providers would like and they have to use 9 of them for something big enough to be interesting. But one of those engines is small enough, just barely, to land a Falcon 9. An Atlas 5 couldn't do it, the one engine's too big. Even with a reusable Falcon 9, it will have to land with >1G of acceleration, it won't be able to hover or descend at a constant rate.

That’s actually not so much of an issue as you might think, provided you design the spacecraft around being able to land propulsively. The answer is throttling, similar to the SSMEs (which can go down to 67% of rated power level); some RL-10s have been modified to throttle down to as little as 5 or 10% of rated thrust. The Falcon 9 has an empty mass of about 20,000 kg as far as I can tell, so a Merlin 1D operating at 10% thrust (ie., producing 65,000 N of thrust at sea level) would accelerate it upwards at only 3.25 m/s^2, rather less than g. No separate landing engines needed.

This was the approach adopted by the DC-X, and would doubtlessly have been adopted by any governmental propulsive vertical landing spacecraft. It makes things more...flexible. It is, however, expensive to design engines to throttle so deeply, and very very expensive and difficult to do it to an existing engine (really, you’ll end up building a new one). So obviously SpaceX isn’t going to do that. If and when they introduce lower stage Raptors (their metholox engines), though, I would be a little bit surprised if they didn’t build one with higher thrust but the ability to throttle relatively deeply.

Megalodon wrote:

I think that was probably doable in the 90s, and we've done much harder things now like self driving cars, but it would have been extremely challenging in the 70s. A winged booster maybe, but that's less mass efficient. Or separate landing engines, but same problem.

Well, almost all of the serious proposals in the 1960s and 1970s were winged, as a matter of course, or at worst lifting bodies. There were a few exceptions, like Phil Bono or the good people at Chrysler Aerospace with SERV (which would have used, from what I recall, around 28 jet engines to land vertically), as well as a few colossal vehicles Boeing, Lockheed, et. al. designed at the end of the 1970s for launching solar power satellite components, but by and large most reusable spacecraft designs have used wings, even through the 1990s up to today. The influence of aeronautical upon aerospace engineering, I suppose, although there are significant merits to winged craft under some circumstances.For instance, unless you’re using what’s known as a pop-up launch trajectory, where the first stage goes almost straight up (very uncommon), the first stage will end up a long ways away from the launch site, meaning that if you want to recover it, whether with parachutes or rocket landings or whatever, you have to have it fly back to where it took off from first. Wings in glideback or flyback mode turn out to be rather handy for this compared to reserving enough propellant to relight the main engines and ballistically boost back, to the point where you can use less mass for a winged than a ballistic vehicle, sometimes much less, depending on the details of your launch trajectory. Of course, sometimes this advantage can be compensated for by other advantages--like being able to reuse designs for an existing vehicle (*cough*Falcon 9). It’s always a process of optimization with a large number of knobs to pull, and a lot of the time options that sound good in the abstract don’t really get you where you want to go.

That’s actually not so much of an issue as you might think, provided you design the spacecraft around being able to land propulsively. The answer is throttling, similar to the SSMEs (which can go down to 67% of rated power level); some RL-10s have been modified to throttle down to as little as 5 or 10% of rated thrust.

The exhaust gasses will expand to ambient pressure before they fill up the nozzle and tear it apart due to flow separation. You can't throttle a rocket engine that low at sea level. An RL-10 might be able to do 5% in vacuum, but that won't be helpful at sea level. Or if its modified to be able to throttle lower, it will be at the expense of expansion ratio, which will harm its performance during the actual flight.

Low throttle isn't an option, and if you look at what SpaceX has said about the latest Grasshopper flight, they specifically mention that it landed with a T/W ratio of >1, and that this is important for their landing algorithm.

truth is life wrote:

If and when they introduce lower stage Raptors (their metholox engines), though, I would be a little bit surprised if they didn’t build one with higher thrust but the ability to throttle relatively deeply.

As a staged combustion engine it may have a higher chamber pressure and be able to throttle lower at sea level as a result, but due to gravity losses landing at higher thrust is optimal. If they've got it working at higher thrust, there's no incentive to go back.

truth is life wrote:

For instance, unless you’re using what’s known as a pop-up launch trajectory, where the first stage goes almost straight up (very uncommon), the first stage will end up a long ways away from the launch site, meaning that if you want to recover it, whether with parachutes or rocket landings or whatever, you have to have it fly back to where it took off from first. Wings in glideback or flyback mode turn out to be rather handy for this compared to reserving enough propellant to relight the main engines and ballistically boost back, to the point where you can use less mass for a winged than a ballistic vehicle, sometimes much less, depending on the details of your launch trajectory. Of course, sometimes this advantage can be compensated for by other advantages--like being able to reuse designs for an existing vehicle (*cough*Falcon 9). It’s always a process of optimization with a large number of knobs to pull, and a lot of the time options that sound good in the abstract don’t really get you where you want to go.

They were already helping Stratolaunch with their airlifted launcher, which includes wings on the first stage. I think we can assume this option has been considered.

The exhaust gasses will expand to ambient pressure before they fill up the nozzle and tear it apart due to flow separation. You can't throttle a rocket engine that low at sea level. An RL-10 might be able to do 5% in vacuum, but that won't be helpful at sea level. Or if its modified to be able to throttle lower, it will be at the expense of expansion ratio, which will harm its performance during the actual flight.

2: You obviously don’t need to throttle so low to begin with, given that a 10% throttled Merlin 1D produced a T/W ratio of about 0.2. Merely throttling to, say, 40% would put you pretty close to 1 in T/W ratio at empty and is achievable if difficult in atmosphere (as shown by the RL10-A-5 used on the DC-X).

3: This is an example of the tradeoffs I’ve been mentioning, and returns to the fact that you can overoptimize in a particular direction. If you are designing from scratch a reusable VTVL rocket, it’s generally desirable to have it be able to produce a T/W < 1 at empty to provide landing margin; this requires either dedicated landing rockets or throttling issues. Both have their problems; this is one of them.

Regardless, you are going to take performance penalties for going reusable...the choice is which ones you’re going to take and why. Sacrificing some of your ISP and thrust to achieve deeper throttability and reuse of existing engines as landing engines may be one of those, depending on the specific design.

Megalodon wrote:

Low throttle isn't an option, and if you look at what SpaceX has said about the latest Grasshopper flight, they specifically mention that it landed with a T/W ratio of >1, and that this is important for their landing algorithm.

Well, of course it’s not an option. The Merlin is probably designed to throttle to avoid overstressing during max-Q, but I would be highly surprised if they had designed it from the outset to throttle down to, say, 35 or 40% to function as a landing engine. That would be expensive and complicated to do, which would have been highly undesirable while it was originally being designed in 2003 and were planning on parachute recovery. And when they much later decided that powered landing was superior, it would have been a difficult and, again, expensive job to add throttling capability into the engines. So they were forced to compensate for physical limitations through more sophisticated software to handle T/W > 1.

The question of whether to add throttling is only relevant for a new clean-slate engine, which brings us to...

Megalodon wrote:

As a staged combustion engine it may have a higher chamber pressure and be able to throttle lower at sea level as a result, but due to gravity losses landing at higher thrust is optimal. If they've got it working at higher thrust, there's no incentive to go back.

Higher thrust is better while you’re decelerating to a stop, true. It most certainly is not while you’re doing terminal maneuvers, because it removes a large amount of margin for correcting errors during earlier operations. You have to get it right the first time, or you’re going to crash. Now, with modern GPS and software, it’s pretty likely you’re going to get it right the first time...but still not 100% certain.

Megalodon wrote:

They were already helping Stratolaunch with their airlifted launcher, which includes wings on the first stage. I think we can assume this option has been considered.

No, actually we can’t. Look at Pegasus, the much smaller distant relative of Falcon 9 Air: it, too, has wings on the first stage, but it would be absurd to suggest that it could be a winged reusable launch vehicle because of that. The airlaunch and surface recovery environments are significantly different, and a great deal of supplemental equipment and capabilities are required for the latter which are not for the former. Fundamentally, conventional rockets like the Falcon 9 are not suitable for winged (much less powered) recovery without so much modification and redesign that you might as well design a new launch vehicle designed for the capability from the outset.

Well, of course it’s not an option. The Merlin is probably designed to throttle to avoid overstressing during max-Q

It hasn't supported throttling at all until the current version. They actually have to shut down 2 of the engines shortly before staging because the nearly empty stage in vacuum would otherwise exceed acceptable acceleration, and they can't do anything at Max-Q.

The new engine is the first one that can throttle. It hasn't flown yet on an orbital launch, it's only flown at all on Grasshopper hops.

(N.B.: the Merlin Vacuum has always been able to throttle, but the first stage engines never have)

truth is life wrote:

Higher thrust is better while you’re decelerating to a stop, true. It most certainly is not while you’re doing terminal maneuvers, because it removes a large amount of margin for correcting errors during earlier operations. You have to get it right the first time, or you’re going to crash.

Even if you could throttle low enough you'd need fuel margin they aren't willing to give it.